Thermal protection of gas turbine combustors relies heavily upon the delivery of a carefully managed film of coolant air to the hot-side of the combustor liner. Furthermore, improvements in engine sfc and the trend to ever more aggressive engine cycles means greater emphasis is being placed upon more efficient use of the proportion of combustion system air made available for cooling. As a result, there is a requirement to better understand the development of cooling films deposited onto the hot-side of the liner through complex effusion arrays. This study, therefore, is concerned with the prediction and measurement of adiabatic film effectiveness of a number of engine-representative designs. A RANS based CFD approach is used to predict film effectiveness in which computational cost is minimised by solving first for a single coolant passage to provide high fidelity, near-exit boundary conditions to the effusion arrays. Equivalent measurements are made for each test case using a Pressure Sensitive Paint (PSP) technique in which the oxygen-quenched fluorescence properties of the paint are employed together with a Nitrogen gas cooling simulant to determine adiabatic film effectiveness. This study demonstrates that whist the model under-predicts the mixing of the coolant with the main-stream flow, and hence the film development over the surface, the approach works well at quantifying the relative performance of each design.

Full-coverage or effusion cooling is commonly used in the thermal management of gas turbine combustion systems. The combustor environment is characterised by highly turbulent freestream flow conditions and relatively large turbulent length scales (length scale-to-coolant hole diameter ratios in excess of 30) that are primarily created by the fuel injector and dilution jets; indeed, the available evidence suggests that large energetic eddies interact strongly with the coolant flows and may have a significant impact on the film-cooling performance. The desire to create compact low-emission combustion systems for aero gas turbine engines has also given rise to a desire to reduce the quantity of air used in wall cooling, and has led to the need for improved thermal design approaches, cooling correlations and validated computational methods. In order to establish a greater understanding of effusion cooling under conditions of very high freestream turbulence, a new laboratory-based test facility has been created that is capable of simulating representative combustor flow conditions, and that allows for a systematic investigation of film-cooling performance over a range of freestream turbulence conditions and coolant to mainstream density ratios. This paper describes the new test facility and its capabilities, including the method for generating combustor relevant flow conditions. Adiabatic film-cooling effectiveness data obtained at a range of blowing ratios are also presented for a typical combustor effusion cooling geometry that uses a twenty degree injection angle. The analysis of this data is supported by fluid temperature field measurements which are presented for low and high freestream turbulence conditions. The interpretation of the data has established the impact of turbulence intensity and integral length scale on the mixing processes between freestream and coolant flows.

The requirements to reduce engine fuel burn costs and gaseous emissions combine to ensure that gas turbine engine manufacturers continually seek to increase the peak cycle temperatures of new engine designs. Consequently, high-pressure turbine components must be developed that can withstand increasing gas temperatures, resulting in the continuous introduction of new technologies that allow appropriate service life. Accurate gas path measurements are vital for early understanding of the performance of a new design, although the accurate measurement of fluid temperature in a turbine stage is becoming increasingly difficult. The use of active probe-assembly cooling is important to ensure a sufficiently robust measurement system. Cooling issues may also affect the measurement performance because the component to which the temperature instrumentation is attached is cooled (for example, a guide vane). The use of a radial traverse total temperature device in the turbine section of a large civil aero-engine has previously been reported and the results analysed and compared to expectation. One outcome of ongoing work has been a proposal for a new design of turbine traverse probe with improved total temperature measurement accuracy. The new design directly addresses those uncertainties caused by conduction of heat from the thermocouple junction and into the cooled probe support. Extensive conjugate CFD modelling followed by validation tests in a high temperature free-jet rig confirmed the success of the design in reducing the magnitude of the thermal conduction error. The probe is likely to be used in future tests to improve engine performance validation.

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